The disclosed embodiments generally pertain to a non-metallic booster spool for use in a gas turbine engine. More particularly, but not by way of limitation, present embodiments relate to a non-metallic, for example composite, booster spool with separable non-metallic, for example composite, blades extending through the booster spool.
A typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween. An air inlet or intake is at a forward end of the engine. Moving toward the aft end, in order, the intake is followed by an engine core including a high pressure compressor, a combustion chamber, and a high pressure turbine. It will be readily apparent from those skilled in the art that additional components may also be included in the engine, such as, for example, low-pressure compressors and low-pressure turbines. This, however, is not an exhaustive list. An engine also typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The internal shaft is connected to both the turbine and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades.
In operation, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. In a two stage turbine, a second stage stator nozzle assembly is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second supporting rotor disk. The turbine converts the combustion gas energy to mechanical energy wherein each set of stator vanes turns and accelerates the combustion gases to engage an adjacent row of rotating turbine blades.
Located aft of the fan is a low pressure compressor, also referred to as a booster. The booster comprises a spool which rotates with a plurality of blades to increase air pressure. Current spool structures are formed of metal or composite/metallic systems. These systems also include blades which are integrally formed with the spool. Such integral formation results in more difficult maintenance for the engine. For example, if one integral blade breaks, the entire spool or large sections thereof must be replaced in order to replace the blade and this results in maintenance which is more difficult and more complex to perform.
As may be seen by the foregoing, it would be desirable to overcome these and other deficiencies within a gas turbine engine. One of the continuous improvement goals in gas turbine engines, especially those used in the aviation industry, is to improve performance through weight reduction of engine components. Accordingly, it would be desirable to provide a more lightweight assembly in the low pressure compressor, or booster, to decrease weight of the assembly and improve performance. It would also be desirable to provide such lightweight configuration in a manner that need not require replacement of entire assemblies or large portions thereof for a single structure breakage or replacement.
The information included in this Background section of the specification, including any references cited herein and any description or discussion thereof, is included for technical reference purposes only and is not to be regarded subject matter by which the scope of the invention is to be bound.